Rotor blade an axial-flow engine

ABSTRACT

This invention relates to a rotor blade airfoil of an axial-flow turbomachine, more particularly for a fan or a compressor high-pressure stage of a gas turbine system, having means in the tip area to improve supersonic performance. In accordance with the present invention, the blade afflux edge has a forward-back sweep. The tip region of the blade efflux edge can be designed essentially similar to the blade afflux edge, with the tip region extending across the outer 15% to 25% of the blade span.

[0001] This invention relates to a rotor blade airfoil of an axial-flow turbomachine, more particularly for a fan or a compressor high-pressure stage of a gas turbine system, having means in the tip area to improve supersonic performance. Associated state of the art is described in U.S. Pat. No. 3,989,406 and U.S. Pat. No. 4,012,172. The first-cited patent specification deals with the supersonic performance of rotor blade airfoils generally, while the second illustrates a forward swept airfoil, i.e. an airfoil on which the leading edge, starting from the blade root, first arches counter to the direction of flow, so that in a lateral view, the airfoil area or blade chord first continuously increases until, upon reaching a radius point, it arches in the direction of flow toward the blade tip area, where in said lateral view, the airfoil area or chord of the blade again decreases.

[0002] The present invention also relates to the sweep of rotor blade airfoils, more particularly of the first compressor high-pressure stages or of a fan of gas turbines or aircraft gas turbine engines. Owing to the high peripheral speeds of said components and the practically zero-swirl inlet flow to the blade, supersonic regions occur in the blade tip area with mach numbers in excess of 1.4. These for several reasons negatively affect the performance of this component. Firstly, the efficiency decreases with increasing blade span, owing to growing shock losses, much more severely than with subsonic rotors. Also, the interaction of duct shock with blade tip swirl negatively affects the stability of the flow, because large blockage regions occur in that area whose nonlinear growth ultimately determines the stability limit of the compressor.

[0003] In a broad aspect, the arrangement of the present invention provides means to remedy said problems.

[0004] It is a particular object of the present invention to provide a solution to said problems by providing a forward-back-swept blade leading edge. Further objects and advantages of the present invention will become apparent from the subclaims.

[0005] The invention is described more fully by means of a preferred embodiment in the light of a single drawing showing a rotor blade airfoil in greatly simplified lateral view.

[0006] Reference numeral 1 indicates the blade root and reference numeral 2 the airfoil of the rotor blade of an axial-flow turbomachine, more particularly for a fan or compressor high-pressure stage of a gas turbine system. Air is impinging on airfoil 2 in the direction of arrowhead 3 (=direction of flow 3) , i.e. the left-hand leading edge of airfoil 2 is its afflux edge 4 a and the right-hand trailing edge is its efflux edge 4 b. The axis of rotation of the rotor, omitted on the drawing, accordingly runs parallel to the direction of flow 3 and extends (far) below the blade root 1.

[0007] As usual, airfoil 2 has a root region 2 a and a tip region 2 b, the latter substantially extending along the outer 15% to 25%, more particularly along about 20% of the blade span h. The latter by amount is the difference (R_(a)−R_(i)), where R_(a) is the outer radius measured from the axis of rotation and R_(i) the inner radius of the airfoil 2. Shown in the tip region 2 b is both the afflux edge 4 a and the efflux edge 4 b of a conventionally designed airfoil 2, indicated by dashed lines a′ and b′. A conventionally shaped airfoil of this type, however, is impaired by the disadvantages cited above.

[0008] To remedy these disadvantages, the airfoil 2 of the present invention is in the tip region 2 b provided with a forward-back sweep at least on the blade afflux edge 4 a as shown on the drawing, i.e. in this tip region 2 b the blade afflux edge 4 a extends, unlike in a conventional design, first counter to the direction of flow 3 and, upon reaching a radius point U, retreats in the direction of flow 3 in a manner more pronounced than with a conventionally designed blade afflux edge a′. For mechanical reasons, an angle α of a 90° order of magnitude is sought in the uppermost region, i.e. at the blade tip, between the afflux edge 4 a and the case contour indicated by line 5 of a case (omitted on the drawing) surrounding the rotor.

[0009] In the tip region 2 b, the blade efflux edge 4 b has a shape essentially similar to that of the blade afflux edge 4 a, i.e. it reflects the forward-back sweep, as a comparison of the solid line 4 b with the line b′ representing the conventional design will show.

[0010] In this arrangement, the back sweep can be made more pronounced than the forward sweep, so as to allow for the mach number distribution. This means that the distance, not indicated on the drawing, by which the blade afflux edge 4 a in the tip region is shifted back compared with the conventional design in accordance with line a′, is larger than the distance, again not indicated, by which the blade afflux edge 4 a in the radius point U area is shifted forward relative to the conventional design in accordance with line a′. On the drawing, this may look different on account of the two-dimensional representation, whereas a rotor blade airfoil is obviously a three-dimensional shape.

[0011] The sweep can also be designed such that the topmost airfoil section of a conventional rotor blade airfoil is retained. This will allow the blading in a previously existing case of an axial-flow turbomachine also to be renewed in a case with contoured walls, without having to adjust rub rings.

[0012] The forward-back sweep of the present invention primarily reduces the local afflux mach number and so reduces shock losses. In the process, the forward-back sweep does not induce greater radial velocities that would be the cause of additional losses.

[0013] A forward-back sweep also provides structural mechanical advantages, considering that the thrust, elastic axis and gravity centers will, when compared with the straight forward or back sweep, not change at all across blade span h until about 80% of the blade span, and only slightly from 80% blade span to the blade tip. Additional bending moments in the hub area, such as arising with straight forward or back sweep, are therefore avoided. The modified modes of vibration in the upper region of the airfoil 2 are above the second E.O., so that the arrangement of the present invention also causes no restrictions from the structural dynamic aspect. It is apparent that a plurality of especially design features other than those described herein may be incorporated in the present embodiment without departing from the inventive concept.

LIST OF REFERENCES

[0014]1 Blade root

[0015]2 Airfoil

[0016]2 a Root region of 2

[0017]2 b Tip region of 2

[0018]3 Direction of flow

[0019]4 a Blade afflux edge

[0020]4 b Blade efflux edge

[0021]5 Case contour

[0022] R_(a) Outer radius of 2

[0023] R_(i) Inner radius of 2

[0024] W Radius point

[0025] a′ Conventional contour of 4 a in 2 b

[0026] b′ Conventional contour of 4 b in 2 b

[0027] h Blade span

[0028] α Angle between 4 a and 5 

What is claimed is:
 1. Rotor blade airfoil (2) of an axial-flow turbomachine, more particularly for a fan or a compressor high-pressure stage of a gas turbine system, having means in the tip region (2 b) for improving supersonic performance, characterized by a forward-back sweep of the blade afflux edge 4 a):
 2. Rotor blade airfoil of claim 1 , characterized in that the tip region (2 b) of the blade efflux edge (4 b) is designed essentially similar to that of the blade afflux edge (4 a).
 3. Rotor blade of claim 1 or 2 , characterized in that the tip region (2 b) extends across the outer 15% to 25% of the blade span (h). 